As is known, a twin-engine or three-engine helicopter has a propulsion system comprising two or three turboshaft engines, each turboshaft engine comprising a gas generator and a free turbine which is rotated by the gas generator and is rigidly connected to an output shaft. The output shaft of each free turbine is suitable for putting into motion a power transmission gearbox, which itself drives the rotor of the helicopter.
It is known that, when the helicopter is in a cruising flight situation (i.e. when it is progressing in normal conditions, during all the flight phases apart from transitional phases of take-off, climbing, landing or hovering flight), the turboshaft engines operate at low power levels, below their maximum continuous output. These low power levels give rise to a specific consumption (hereinafter Cs), defined as the ratio between the hourly consumption of fuel by the combustion chamber of the turboshaft engine and the mechanical power supplied by this turboshaft engine, greater than approximately 30% of the Cs of the maximum take-off power, and therefore give rise to overconsumption of fuel in cruising flight.
Moreover, the turboshaft engines of a helicopter are designed so as to be oversized in order to be able to keep the helicopter in flight in the event of failure of one of the engines. This flight situation occurs following the loss of an engine, and results in the fact that each functioning engine provides a power that is significantly greater than its rated power in order to allow the helicopter to overcome a dangerous situation, and to then continue its flight.
The turboshaft engines are also oversized so as to be able to ensure flight over the entire flight range specified by the aircraft manufacturer, and in particular flight at high altitudes and during hot weather. These flight points, which are very restrictive, in particular when the helicopter has a weight that is close to its maximum take-off weight, are only encountered in specific use cases.
These oversized turboshaft engines are disadvantageous in terms of weight and fuel consumption. In order to reduce this consumption in cruising flight, it is conceivable to put at least one of the turboshaft engines on standby in flight. The active engine or engines then operate at higher power levels in order to provide all the necessary power, and therefore at more favourable Cs levels.
Putting a turboshaft engine on standby involves a rapid reactivation system that makes it possible to rapidly bring the turboshaft engine out of standby when it is needed. This need may for example arise from one of the active engines failing or from the flight conditions unexpectedly deteriorating, requiring full power to be reinstated rapidly.
The applicant has already proposed a rapid reactivation system that uses a pneumatic turbine that is mechanically connected to the turboshaft engine and is configured such that it can transform the power from the pressurised gas at the turbine inlet into mechanical power that drives the gas generator of the turboshaft engine. The supply of gas to the pneumatic turbine may for example be achieved by the cooperation of a pneumatic store and a controlled fast-opening valve or by a solid-propellant storage device.
The applicant has therefore sought to develop a method and a device for integrity testing the rapid reactivation system such that it is possible to ensure that the rapid reactivation system operates and can be used during flight.